This invention relates generally to gas turbine engines and more particularly to improved airfoil construction. An improved blade construction embodying a member of a first material such as ceramic and shaped as an airfoil cooperatively arranged with a second material such as metal for the insertion in the air flow path of a reduction gas turbine engine. The improved blade can absorb and function with thermally induced stress. Further the enhanced airfoil exhibits durability characteristics suitable for use in a high-temperature environment.
In a well-known gas turbine engine, air is compressed in a rotating compressor, heated in a combustion chamber and expanded through a turbine. To increase the available energy in the turbojet cycle, and hence the thrust and efficiency of the engine, designers have conventionally attempted to increase the turbine inlet temperature because turbine engine power is directly related to turbine inlet temperature, as is turbine engine efficiency.
However, as a result of the high temperatures necessary to operate the gas turbine engines, trailing and leading edges of the turbine's stator blade are often burned away. Repairs to restore the turbine stator blade to functional condition are expensive and require extensive hours of welding. The erosion and burning of the high pressure turbine stator blade is a primary repair frequency item for turbomachinery applications. To improve the life of turbine stator blades, industry has turned to methods such as cooling blades while in use.
One method, therefore, of providing more efficient turbojet operation, i.e. higher cycle temperatures and hence higher thrust values for a given size engine, involves the use of more sophisticated and advanced turbine airfoil cooling techniques to permit higher turbine inlet temperatures. With these techniques, turbine nozzle (vane) and rotor blade temperatures may be brought within the capability of existent heat or oxidation resistant materials (metal). However, there is a limit to temperature that can be achieved with these materials.
As an alternative to such cooling methods or techniques recourse is made to improve blade or vane materials and construction methods. It is known, for example, that ceramic materials have the ability to withstand significantly higher temperatures than the known refractory alloys when used in the conventional blade or vane designs, even when the latter are intensely cooled. In particular, the use of ceramic material would appear to be indicated at the leading edge of the turbine airfoil where the temperatures are always highest and where cooling is most difficult since the heat input is highest here also. It will be understood that the term "ceramic" as herein used includes composites of nonmetallics and metallics, the latter sometimes being referred to as ceramics. Examples of the latter materials which have been evaluated and been found to perform well in the practice of the present invention include: (1) chromium and 30 weight percent M.sub.g O, or Cr30M.sub.g O, and (2)Al.sub.2 O.sub.3 or (3) SI.sub.3 N.sub.4.
However, ceramic materials present certain problems that have heretofore prohibited their widespread use for constructing part or all of the blade or vane. Firstly, ceramic materials do not have the tensile strength of metallic materials. Secondly, due to its usually relatively low ductility, ceramic material has a tendency to crack under the impact of severe or suddenly applied thermal shock or stresses such as may occur in advanced lightweight aircraft gas turbines. Thirdly, where it is desired to strengthen the blade or vane structurally, such as by means of metallic body or strut member or members, the dissimilar characteristics of the metal body and the ceramic materials in the areas of ductility, thermal conductivity and brittleness, for example, create additional problems concerning how to mate these materials in an integral airfoil construction.